The turbine section of a gas turbine engine is subjected to extremely high temperatures. The temperature of hot gases entering the turbine from the combustor is generally well above the melting point temperatures of the alloys from which turbine rotor blades and stator vanes are fabricated. Since both vanes and blades are subjected to such high temperatures, they must be cooled to maintain their structural integrity.
The turbine vanes and blades are cooled by air bled from the engine's compressor, bypassing the combustor. It will be understood that any compressor bleed air for such cooling will be unavailable to support combustion in the combustor. Thus, to minimize any sacrifice in engine performance due to inadequate airflow to support combustion, any scheme for cooling turbine blades and vanes must optimize the utilization of compressor bleed cooling air.
Generally, turbine blade and vane cooling is accomplished by external film cooling, and internal air impingement and convection cooling, or a combination of both. With air impingement, compressor bleed air is channeled to the inside of an airfoil and directed onto the inside walls of the airfoil. The air then exits the airfoil through a set of film holes provided within the airfoil walls. Air impingement is an effective method for cooling blades.
In convection cooling, compressor bleed air flows through typically serpentine passages in the blades and vanes, continuously removing heat therefrom. Compressor bleed air enters the passages through an inlet disposed generally in a leading portion of the airfoil, which discharges into the serpentine passage. The passage also includes fins or ridges (also known as "trip strips") in a wall thereof, which facilitate improved convection cooling of the airfoil walls. The precise dimensions and locations of these trip strips can control the amount of air flow through the passage and also, at least in part, determine the cooling efficiency of the overall serpentine structure.
Turning vanes may be employed to channel air flow around the bends of the serpentine passages which terminate in a series of film holes located near the trailing edge of the blade or vane, through which some of the cooling air is discharged.
It will be appreciated that as the cooling air flows through the serpentine passage, it loses pressure. By the time the air flow reaches the trailing edge, if the pressure is too low, hot combustion gases outside the blade or vane may enter through the blade or vane film holes therein. This would, of course, hinder the cooling of the airfoil and may cause burning of the airfoil walls. Thus, there is an acute need to maintain a higher level of pressure within the airfoil cooling passages than on the outside of the airfoil.
The conventional method of fabricating the airfoil blade or vane with serpentine cooling passages is to cast the part and then to precisely shape the part through extensive machining. In the casting process, a mold and a core are first fabricated, the core defining the shape of the cooling air passages in the interior of the airfoil. The core is held in place by a core support rod attached to the core and protruding through the airfoil wall.
After the casting process is completed, the core is dissolved by application of a chemical solution, the voids in the airfoil left by dissolving the core forming the serpentine cooling passages in the airfoil. The hole within the airfoil casting formed by removal of the core support rod is then closed and the cast part is machined into its final shape.
Although the casting process yields a high quality product, the process itself is costly and time consuming. The fabrication of the mold and core are especially complex due to the intricate shapes of the cooling passages. Thus, retooling to correct airfoil overheating problems discovered during the testing may be expensive and impractical, if not economically prohibitive.
Presently, it is not conveniently possible to adjust the amount of air flow through airfoil cooling passages without the extreme costs and expenditures of time required in retooling to recast cooling passage with adjustments in the geometrics thereof.
Another problem that arises with serpentine cooling passages, is that once the compressor bleed air enters the serpentine passage, it gets progressively hotter within the passage as it removes heat from the airfoil and may not have adequate cooling capacity to effectively cool the trailing edge of the airfoil.
It will be appreciated that any solutions to the problems of cooling current turbine blades and vanes outlined hereinabove, should not come at the expense of added weight, which even in minute quantities thereof per airfoil, can significantly detract from engine performance.